Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

ONERA HOR07 AIRFOIL (hor07-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: ONERA HOR07 AIRFOIL (hor07-il)
Reynolds number: 200,000
Max Cl/Cd: 78.35 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hor07-il-200000.txt
Download as CSV file: xf-hor07-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: ONERA HOR07 AIRFOIL                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.4037   0.09796   0.09461  -0.0056   1.0000   0.0404
  -7.500  -0.3973   0.09488   0.09156  -0.0072   1.0000   0.0415
  -7.250  -0.3912   0.09180   0.08851  -0.0093   1.0000   0.0428
  -7.000  -0.3810   0.08837   0.08511  -0.0135   1.0000   0.0445
  -6.750  -0.3581   0.08446   0.08121  -0.0275   1.0000   0.0465
  -6.500  -0.3333   0.07999   0.07671  -0.0367   1.0000   0.0468
  -6.250  -0.3291   0.07429   0.07108  -0.0359   1.0000   0.0478
  -6.000  -0.3175   0.07150   0.06832  -0.0345   1.0000   0.0490
  -5.750  -0.3011   0.06859   0.06544  -0.0358   1.0000   0.0506
  -5.500  -0.2806   0.06521   0.06207  -0.0395   1.0000   0.0530
  -5.250  -0.2205   0.05839   0.05499  -0.0605   1.0000   0.0591
  -5.000  -0.2108   0.05370   0.05042  -0.0606   0.9953   0.0607
  -4.750  -0.1801   0.05159   0.04831  -0.0623   0.9840   0.0647
  -4.500  -0.1200   0.04435   0.04076  -0.0770   0.9721   0.0742
  -4.250  -0.0858   0.04180   0.03817  -0.0800   0.9566   0.0777
  -4.000  -0.0067   0.02318   0.01789  -0.0980   0.9508   0.0431
  -3.750   0.0336   0.01897   0.01279  -0.1009   0.9342   0.0406
  -3.500   0.0651   0.01720   0.01055  -0.1013   0.9151   0.0414
  -3.250   0.0929   0.01616   0.00919  -0.1008   0.8970   0.0438
  -3.000   0.1198   0.01533   0.00804  -0.1000   0.8804   0.0476
  -2.750   0.1470   0.01424   0.00685  -0.0995   0.8651   0.0528
  -2.250   0.2024   0.01304   0.00546  -0.0987   0.8364   0.0802
  -2.000   0.2305   0.01276   0.00519  -0.0984   0.8227   0.1105
  -1.750   0.2585   0.01259   0.00503  -0.0982   0.8093   0.1575
  -1.500   0.2861   0.01260   0.00497  -0.0979   0.7963   0.1878
  -1.250   0.3137   0.01255   0.00491  -0.0977   0.7841   0.2091
  -1.000   0.3414   0.01251   0.00480  -0.0974   0.7730   0.2284
  -0.750   0.3693   0.01244   0.00470  -0.0972   0.7622   0.2498
  -0.500   0.3975   0.01237   0.00465  -0.0972   0.7511   0.2754
  -0.250   0.4257   0.01224   0.00457  -0.0971   0.7411   0.2992
   0.000   0.4540   0.01208   0.00445  -0.0971   0.7316   0.3319
   0.250   0.4825   0.01183   0.00444  -0.0972   0.7213   0.3876
   0.500   0.5024   0.01028   0.00423  -0.0951   0.7126   1.0000
   0.750   0.5311   0.01041   0.00417  -0.0950   0.7027   1.0000
   1.000   0.5597   0.01053   0.00418  -0.0950   0.6924   1.0000
   1.250   0.5880   0.01067   0.00419  -0.0949   0.6829   1.0000
   1.500   0.6164   0.01079   0.00422  -0.0948   0.6725   1.0000
   1.750   0.6448   0.01090   0.00428  -0.0948   0.6615   1.0000
   2.000   0.6730   0.01103   0.00435  -0.0947   0.6511   1.0000
   2.250   0.7011   0.01116   0.00440  -0.0946   0.6409   1.0000
   2.500   0.7294   0.01128   0.00449  -0.0945   0.6288   1.0000
   2.750   0.7575   0.01142   0.00460  -0.0945   0.6169   1.0000
   3.000   0.7855   0.01157   0.00473  -0.0944   0.6052   1.0000
   3.250   0.8135   0.01174   0.00485  -0.0942   0.5939   1.0000
   3.500   0.8414   0.01191   0.00499  -0.0941   0.5818   1.0000
   3.750   0.8692   0.01207   0.00517  -0.0940   0.5682   1.0000
   4.000   0.8967   0.01221   0.00528  -0.0938   0.5523   1.0000
   4.250   0.9239   0.01228   0.00534  -0.0935   0.5287   1.0000
   4.500   0.9507   0.01241   0.00540  -0.0932   0.5047   1.0000
   4.750   0.9779   0.01260   0.00558  -0.0930   0.4824   1.0000
   5.000   1.0047   0.01285   0.00579  -0.0927   0.4605   1.0000
   5.250   1.0311   0.01316   0.00603  -0.0925   0.4328   1.0000
   5.500   1.0572   0.01351   0.00633  -0.0922   0.3934   1.0000
   5.750   1.0777   0.01494   0.00694  -0.0916   0.2321   1.0000
   6.000   1.0927   0.01797   0.00894  -0.0908   0.0795   1.0000
   6.250   1.1155   0.01909   0.01011  -0.0902   0.0696   1.0000
   6.500   1.1373   0.02025   0.01134  -0.0894   0.0649   1.0000
   6.750   1.1595   0.02125   0.01243  -0.0886   0.0616   1.0000
   7.000   1.1807   0.02236   0.01360  -0.0877   0.0592   1.0000
   7.250   1.2006   0.02365   0.01490  -0.0866   0.0570   1.0000
   7.500   1.2179   0.02569   0.01690  -0.0852   0.0546   1.0000
   7.750   1.2401   0.02683   0.01813  -0.0843   0.0535   1.0000
   8.000   1.2625   0.02812   0.01955  -0.0833   0.0525   1.0000
   8.250   1.2853   0.02958   0.02113  -0.0824   0.0515   1.0000
   8.500   1.3080   0.03106   0.02275  -0.0816   0.0499   1.0000
   8.750   1.3295   0.03236   0.02414  -0.0808   0.0473   1.0000
   9.000   1.3499   0.03416   0.02593  -0.0802   0.0443   1.0000
   9.250   1.3692   0.03670   0.02875  -0.0792   0.0416   1.0000
   9.500   1.3876   0.03813   0.03046  -0.0778   0.0391   1.0000
   9.750   1.4053   0.03999   0.03251  -0.0766   0.0366   1.0000
  10.000   1.4212   0.04203   0.03457  -0.0757   0.0341   1.0000
  10.250   1.4291   0.04598   0.03895  -0.0737   0.0318   1.0000
  10.500   1.4363   0.04853   0.04193  -0.0712   0.0301   1.0000
  10.750   1.4400   0.05198   0.04576  -0.0687   0.0290   1.0000
  11.000   1.4392   0.05558   0.04971  -0.0660   0.0283   1.0000
  11.250   1.4323   0.05936   0.05385  -0.0630   0.0279   1.0000
  11.500   1.4165   0.06317   0.05796  -0.0592   0.0279   1.0000
  11.750   1.3954   0.06740   0.06250  -0.0562   0.0282   1.0000
  12.000   1.3726   0.07232   0.06771  -0.0546   0.0285   1.0000
  12.250   1.3485   0.07785   0.07348  -0.0545   0.0288   1.0000
  12.500   1.3237   0.08393   0.07979  -0.0556   0.0292   1.0000
  12.750   1.2987   0.09055   0.08663  -0.0577   0.0296   1.0000
  13.000   1.2734   0.09777   0.09404  -0.0609   0.0300   1.0000
  13.250   1.2480   0.10566   0.10209  -0.0651   0.0304   1.0000
  13.500   1.2229   0.11424   0.11082  -0.0702   0.0309   1.0000
  13.750   1.1986   0.12354   0.12023  -0.0761   0.0314   1.0000
  14.000   1.1775   0.13304   0.12979  -0.0817   0.0321   1.0000
<< Back to ONERA HOR07 AIRFOIL (hor07-il)

Polar data table (+)

Polar graphs


<< Back to ONERA HOR07 AIRFOIL (hor07-il)